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EPPLER E1212 AIRFOIL (e1212-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER E1212 AIRFOIL (e1212-il)
Reynolds number: 50,000
Max Cl/Cd: 7.07 at α=2.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e1212-il-50000.txt
Download as CSV file: xf-e1212-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER E1212 AIRFOIL                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.000  -0.3088   0.13103   0.12349  -0.0128   1.0000   0.3088
 -10.750  -0.2585   0.12249   0.11488  -0.0122   1.0000   0.3183
 -10.500  -0.2818   0.12372   0.11619  -0.0128   1.0000   0.3283
 -10.250  -0.2492   0.11719   0.10964  -0.0128   1.0000   0.3352
 -10.000  -0.2527   0.11623   0.10874  -0.0128   1.0000   0.3477
  -9.750  -0.2415   0.11231   0.10485  -0.0131   1.0000   0.3540
  -9.500  -0.2296   0.10980   0.10238  -0.0128   1.0000   0.3663
  -9.250  -0.2324   0.10757   0.10022  -0.0129   1.0000   0.3740
  -9.000  -0.2132   0.10442   0.09709  -0.0126   1.0000   0.3860
  -8.750  -0.2197   0.10270   0.09547  -0.0123   1.0000   0.3951
  -8.500  -0.2009   0.09975   0.09256  -0.0119   1.0000   0.4080
  -8.250  -0.2038   0.09776   0.09067  -0.0113   1.0000   0.4171
  -8.000  -0.1947   0.09589   0.08889  -0.0105   1.0000   0.4315
  -7.750  -0.1868   0.09310   0.08621  -0.0099   1.0000   0.4403
  -7.500  -0.2027   0.09343   0.08674  -0.0075   1.0000   0.4550
  -7.250  -0.1840   0.08999   0.08344  -0.0067   1.0000   0.4633
  -7.000  -0.2200   0.09209   0.08583   0.0002   1.0000   0.4682
  -6.750  -0.2823   0.09637   0.09032   0.0045   0.9933   0.4775
  -6.500  -0.2347   0.09211   0.08598  -0.0031   0.9771   0.4991
  -6.250  -0.1752   0.08680   0.08056  -0.0107   0.9613   0.5213
  -6.000  -0.1145   0.08172   0.07538  -0.0184   0.9459   0.5440
  -5.750  -0.0444   0.07642   0.06995  -0.0272   0.9330   0.5693
  -5.500   0.0157   0.07211   0.06553  -0.0350   0.9181   0.5962
  -5.250   0.0367   0.07096   0.06430  -0.0379   0.9001   0.6263
  -5.000  -0.1080   0.06447   0.05770  -0.0479   0.8737   0.4327
  -4.750  -0.1091   0.05895   0.05199  -0.0547   0.8586   0.4072
  -4.500  -0.1065   0.05571   0.04862  -0.0573   0.8421   0.4032
  -4.250  -0.1093   0.05274   0.04548  -0.0592   0.8257   0.4007
  -4.000  -0.0984   0.05050   0.04307  -0.0605   0.8109   0.4033
  -3.750  -0.0749   0.04780   0.04011  -0.0635   0.7991   0.4077
  -3.500  -0.0704   0.04619   0.03827  -0.0638   0.7835   0.4108
  -3.250  -0.0542   0.04489   0.03680  -0.0639   0.7700   0.4157
  -3.000  -0.0220   0.04390   0.03575  -0.0640   0.7585   0.4227
  -2.750  -0.0123   0.04336   0.03508  -0.0632   0.7440   0.4276
  -2.500   0.0143   0.04196   0.03331  -0.0647   0.7335   0.4353
  -2.250   0.0312   0.04182   0.03323  -0.0631   0.7200   0.4408
  -2.000   0.0502   0.04157   0.03289  -0.0624   0.7082   0.4474
  -1.750   0.0758   0.04084   0.03185  -0.0632   0.6973   0.4560
  -1.500   0.0896   0.04108   0.03217  -0.0615   0.6850   0.4621
  -1.250   0.1196   0.04054   0.03150  -0.0615   0.6753   0.4713
  -1.000   0.1304   0.04104   0.03186  -0.0609   0.6632   0.4785
  -0.750   0.1603   0.04066   0.03150  -0.0602   0.6542   0.4884
  -0.500   0.1687   0.04155   0.03230  -0.0593   0.6423   0.4968
  -0.250   0.1985   0.04125   0.03198  -0.0589   0.6336   0.5074
   0.000   0.2053   0.04246   0.03317  -0.0578   0.6228   0.5167
   0.250   0.2321   0.04246   0.03316  -0.0573   0.6138   0.5295
   0.500   0.2426   0.04363   0.03426  -0.0565   0.6043   0.5418
   0.750   0.2540   0.04461   0.03534  -0.0552   0.5950   0.5531
   1.000   0.3001   0.04365   0.03431  -0.0552   0.5888   0.5770
   1.250   0.2696   0.04751   0.03825  -0.0532   0.5772   0.5816
   1.500   0.3126   0.04677   0.03757  -0.0530   0.5706   0.6108
   1.750   0.2864   0.05053   0.04142  -0.0511   0.5621   0.6187
   2.000   0.2857   0.05240   0.04338  -0.0496   0.5546   0.6390
   2.250   0.3528   0.04991   0.04120  -0.0489   0.5485   0.7127
   2.500   0.2905   0.05610   0.04744  -0.0469   0.5432   0.7051
   2.750   0.2780   0.05865   0.05028  -0.0452   0.5386   0.7478
   3.000   0.3813   0.05735   0.04906  -0.0534   0.5269   1.0000
   3.250   0.1888   0.07105   0.06294  -0.0494   0.6170   0.7488
   3.500   0.2524   0.07367   0.06569  -0.0572   0.6108   1.0000
   3.750   0.2813   0.07502   0.06669  -0.0575   0.5900   1.0000
   4.000   0.3329   0.07242   0.06367  -0.0542   0.5311   1.0000
   4.250   0.3613   0.07485   0.06593  -0.0549   0.5274   1.0000
   4.500   0.2460   0.08349   0.07493  -0.0584   0.6355   1.0000
   4.750   0.2803   0.08756   0.07882  -0.0604   0.6299   1.0000
   5.000   0.2701   0.08764   0.07882  -0.0583   0.6169   1.0000
   5.250   0.3116   0.09197   0.08299  -0.0606   0.6106   1.0000
   5.500   0.2957   0.09203   0.08299  -0.0583   0.5984   1.0000
   5.750   0.3331   0.09587   0.08671  -0.0600   0.5906   1.0000
   6.000   0.3221   0.09666   0.08745  -0.0585   0.5807   1.0000
   6.250   0.3450   0.09941   0.09011  -0.0591   0.5725   1.0000
   6.500   0.3772   0.10401   0.09460  -0.0606   0.5681   1.0000
   6.750   0.3624   0.10365   0.09422  -0.0588   0.5551   1.0000
   7.000   0.4017   0.10827   0.09876  -0.0604   0.5491   1.0000
   7.250   0.3838   0.10841   0.09887  -0.0589   0.5391   1.0000
   7.500   0.4056   0.11132   0.10172  -0.0594   0.5319   1.0000
   8.000   0.3420   0.11846   0.10969  -0.0516   0.5061   1.0000
   8.250   0.3388   0.11946   0.11069  -0.0510   0.4949   1.0000
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